1. Field of the Invention
The present invention relates generally to fluid reaction surfaces, and more specifically to turbine airfoils with cooling circuits.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
In a gas turbine engine, especially in an industrial gas turbine engine, a hot gas flow generated in a combustor is passed through a series of rows or stages of turbine stator vanes and rotor blades to convert the thermal energy of the flow into mechanical energy by driving the rotor shaft. The efficiency of the engine can be increased by passing a higher hot gas flow through the turbine. However, the maximum temperature is dependent upon the material properties of the turbine airfoils, especially the first stage vanes and blades because these are exposed to the hottest temperature.
Turbine airfoils can be exposed to higher temperatures than the material properties would allow by passing pressurized cooling air through the airfoils to produce convection cooling, impingement cooling and film cooling of the airfoils. Maximizing the amount of airfoil cooling while minimizing the amount of cooling air used would provide the maximum efficiency for the engine. The rotor blades also are exposed to high gas flow temperatures at the blade tip because of leakage flow. High temperature turbine blade tip section heat load is a function of the blade tip leakage flow. A high leakage flow will induce high heat load onto the blade tip section and therefore the blade tip section sealing and cooling have to be addressed as a single problem. A prior art turbine blade tip includes a squealer tip rail which extends around the perimeter of the airfoil flush with the airfoil wall and forms an inner squealer pocket. FIG. 3 shows a prior art turbine blade with this type of squealer tip design. The main purpose for incorporating a squealer tip in the blade design is to reduce the blade tip leakage and also to provide a rubbing capability for the blade. The squealer tip rail is thin compared to a solid blade tip and therefore rubbing causes less damage to the tip or the shroud surface.
FIG. 2 shows a prior art blade with a squealer tip cooling design. Film cooling holes are built in along the airfoil pressure side tip section. Also, convective cooling holes are positioned along the tip rail at the inner portion of the squealer pocket to provide additional cooling for the squealer tip rail. Secondary hot gas flow migration around the blade tip section is shown by the arrows in FIG. 2.
FIG. 3 shows a prior art blade tip with cooling design for the blade suction side tip rail. The suction side blade tip rail is subject to heating from three exposed sides, and cooling of the suction side squealer tip rail by means of a discharge row of film cooling holes along the pressure side peripheral and at the bottom of the squealer floor becomes insufficient. This is primarily due to the combination of tip rail geometry and the interaction of hot gas secondary flow mixing; the effectiveness induced by the pressure side film cooling and tip section convective cooling holes is very limited.
Turbine blade cooling not only allows for a higher gas flow temperature exposed to the airfoil, but also reduces the occurrence of hot spots around the blade that leads to erosion and spallation, thus shortening the life of the blade.
It is therefore an object of the present invention to provide for a turbine blade with a near wall cooling circuit and a squealer tip cooling design that can be used in a blade cooling design in addition to a passive clearance control system, especially for the blade design with a single suction side tip rail.